CN109960878A - A kind of active/passive thermal protection system coupling design method towards hypersonic aircraft totality - Google Patents
A kind of active/passive thermal protection system coupling design method towards hypersonic aircraft totality Download PDFInfo
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Abstract
The present invention proposes a kind of active/passive thermal protection system coupling design method towards hypersonic aircraft totality, it is first input with aircraft configurations and grid, task trajectory and active cooling parameter, based on equation of heat balance, utilizing works algorithm, which solves, obtains Aerodynamic Heating；Secondly selection thermal protection concept obtains the passive thermal protection concept distribution of entire hypersonic aircraft；It includes the mass flow of cooling working medium and the temperature rise of permission that the scale for determining passive thermal protection system later, which includes the capacity of thickness and quality and active thermal-protection system,；Assessment and Iterative Design are finally carried out until meeting Aircraft Conceptual Design demand.Active Cooling System is placed under passive thermal protection system wall surface by the present invention, to when carrying out the calculating of aircraft surface Aerodynamic Heating, the distribution of passive thermal protection concept and passive thermal protection system scale calculating, the influence for considering active cooling, realizes that aircraft is passive and the Coupling Design of active thermal-protection system.
Description
Technical field
The present invention relates to aircraft thermal protection system design field, more particularly to a kind of towards hypersonic aircraft
Active cooling and passive thermal protection system coupling design method.
Background technique
Currently, with the development of hypersonic engineering, especially aerospace craft is reusable, wide envelope curve flight and long
The appearance of time high-performance cruise demand, aircraft thermal environment very severe, single passive thermal protection technology are unable to satisfy solar heat protection
Demand, it is therefore desirable to which exploitation is passively the same as the thermal protection system actively combined.Passive thermal protection system is usually prevented by different heat
Concept composition is protected, as NASA is based on space shuttle engineering developme a series of metals, ceramic insulation tile and ceramic insulation felt (D E
Myers,C J Martin,and M L Blosser,Parametric Weight Comparison of Advanced
Metallic,Ceramic Tile,and Ceramic Blanket Thermal Protection Systems.2000,
NASA Langley Technical Report Server.), by the anti-of anti-thermal concept, heat-proof quality obstruct Aerodynamic Heating into
Enter body；And active thermal-protection system can be brought heat to low temperature from high-temperature area by cooling down forming to flow network etc. for working medium
Region, to protect aircraft security.Active cooling will affect the Aerodynamic Heating of aircraft wall surface, and then it is general to influence passive thermal protection
Distribution, size and the quality of thought, but in current research, the usual degree of coupling of design actively and passively between thermal protection system is very
It is low, it is difficult to obtain the optimal design section of aircraft thermal protection system.
Summary of the invention
In order to solve the problems existing in the prior art, the present invention propose a kind of active towards hypersonic aircraft totality with
Active Cooling System is placed in passively by the coupling design method of passive thermal protection system by establishing equivalent thermal balance model
Under thermal protection system wall surface, and face is realized by Iterative Design using the intensity of equivalent heat transfer coefficient characterization active cooling
The Coupling Design of the passive and active thermal-protection system overall to hypersonic aircraft.
The technical solution of the present invention is as follows:
A kind of active/passive thermal protection system coupling design method towards hypersonic aircraft totality, feature
It is: the following steps are included:
Step 1: being defeated with the 3-d modelling of hypersonic aircraft and grid division, task trajectory and active cooling parameter
Enter parameter, solved based on equation of heat balance utilizing works algorithm and obtain hypersonic aircraft Aerodynamic Heating, exports aero-thermal load, wall
Face temperature and active heat exchange amount；The active cooling parameter includes equivalent heat transfer coefficient, cooling Temperature of Working；
Step 2: according to the wall surface temperature obtained in step 1, determining maximum wall surface temperature of the analyzed area under overall trajectory；
Using maximum wall surface temperature of the analyzed area under overall trajectory as input condition, it is from the thermal protection conceptual database established in advance
The analyzed area of hypersonic aircraft wall surface selects thermal protection concept, it is desirable that maximum wall surface of the analyzed area under overall trajectory
Temperature is less than the allowable temperature of selected thermal protection concept, and then obtains the passive thermal protection concept of entire hypersonic aircraft
Distribution；
Step 3: establishing one-dimensional and unsteady state heat transfer numerical model for the passive thermal protection concept that step 2 determines, and with step
Rapid 1 obtained aero-thermal load and active heat exchange amount are respectively outer, internal boundary condition progress analysis of Heat Transfer, and it is general to calculate passive thermal protection
The Temperature Distribution of thought；It is then constraint with the set temperature of inner boundary and the allowable temperature of passive thermal protection concept layers of material,
With passive thermal protection system quality most gently for target, in passive thermal protection concept thermal insulation layer with a thickness of design parameter, carry out
Structure optimization calculates the thickness and quality for obtaining optimal passive thermal protection system；It is obtained with cooling down specific heat and the step 1 of working medium
The active heat exchange amount arrived is input, calculates the temperature rise of the mass flow and permission that obtain cooling working medium,
Step 4: needed for the quality and thickness, active thermal-protection system of the passive thermal protection system that judgment step 3 obtains
Whether cooling working medium mass flow meets hypersonic aircraft overall design constraints, if it is satisfied, then current system can be used as
The active/passive thermal protection system of aircraft；If not satisfied, then modifying active cooling parameter, step 1 is repeated to 3, carries out a new round
Design, such iteration is until finally meet hypersonic aircraft overall design constraints.
Further preferred embodiment, a kind of active/passive thermal protection system coupling towards hypersonic aircraft totality
Close design method, it is characterised in that: task trajectory includes flight time, height, speed, the angle of attack and yaw angle in step 1.
Further preferred embodiment, a kind of active/passive thermal protection system coupling towards hypersonic aircraft totality
Close design method, it is characterised in that: equation of heat balance described in step 1 is that Aerodynamic Heating amount is equal to radiation dissipation amount and actively changes
The sum of heat；The active heat exchange amount is equal to wall surface temperature and cooling working medium temperature difference multiplied by equivalent heat transfer coefficient.
Further preferred embodiment, a kind of active/passive thermal protection system coupling towards hypersonic aircraft totality
Close design method, it is characterised in that: the Engineering Algorithm used in step 1 is based on prandtl boundary layer theory, is regarded as outside boundary layer
Perfect gas has viscous gas in boundary layer, by boundary layer outside calculating obtain aerodynamic parameter outside boundary layer, utilize ginseng later
It examines temperature method and calculates the Aerodynamic Heating obtained in boundary layer.
Further preferred embodiment, a kind of active/passive thermal protection system coupling towards hypersonic aircraft totality
Close design method, it is characterised in that: thermal protection conceptual database includes that the density of each layer of thermal protection concept, maximum are permitted in step 2
With temperature, radiant emissivity, thermal coefficient, specific heat and original depth.
Beneficial effect
The present invention is distributed determining, passive thermal protection system matter in progress flight vehicle aerodynamic heat calculating, passive thermal protection concept
Amount and when THICKNESS CALCULATION, coupling considers the influence of active cooling, and working medium needed for can exporting simultaneously active thermal-protection system
Mass flow and working medium temperature rise, realize the coupled relation passively between active thermal-protection system.
Additional aspect and advantage of the invention will be set forth in part in the description, and will partially become from the following description
Obviously, or practice through the invention is recognized.
Detailed description of the invention
Above-mentioned and/or additional aspect of the invention and advantage will become from the description of the embodiment in conjunction with the following figures
Obviously and it is readily appreciated that, in which:
Fig. 1 is active/passive thermal protection system Coupling Design process.
Fig. 2 is thermal protection system thermal balance schematic diagram.
Fig. 3 is equivalent thermal protection system thermal balance schematic diagram.
Fig. 4 is one-dimensional plate finite difference simulator.
Specific embodiment
When the passive thermal protection system of aircraft designs both at home and abroad at present, the influence for considering active cooling, this hair can not be coupled
It is bright to be directed to this problem, propose it is a kind of passively with the coupling design method of active thermal-protection system.It is actively cold in Practical Project
But it is located at the inner boundary bottom of passive thermal protection system, the present invention establishes equivalent thermal balance model, i.e., sets Active Cooling System
Under passive thermal protection system wall surface, thus carry out the calculating of aircraft surface Aerodynamic Heating, passive thermal protection concept distribution with
And passive thermal protection system scale is when calculating, it is contemplated that the influence of active cooling realizes that aircraft is passive and active thermal protection
The Coupling Design of system.
Technical solution of the present invention includes four steps, as shown in Figure 1:
The first step is to calculate Aerodynamic Heating.With aircraft configurations and grid, task trajectory include the flight time, height, speed,
The angle of attack and yaw angle and active cooling parameter include equivalent heat transfer coefficient, cooling Temperature of Working, to input parameter, based on heat
Equilibrium equation, i.e. Aerodynamic Heating amount are equal to the sum of radiation dissipation amount and active heat exchange amount, and utilizing works algorithm, which solves, to be obtained pneumatically
Heat.In the present invention, active heat exchange amount is equal to wall surface temperature and cooling working medium temperature difference multiplied by equivalent heat transfer coefficient, and practical heat is flat
Weighing apparatus model and equivalent thermal balance model are shown in Fig. 2 and Fig. 3 respectively.Engineering Algorithm is based on prandtl boundary layer theory, side in the present invention
Be regarded as perfect gas outside interlayer, to there is viscous gas in boundary layer, by boundary layer outside calculating acquisition boundary layer outer rim pneumatically join
Number, and then the Aerodynamic Heating obtained in boundary layer is calculated using reference temperature method.This step will export aero-thermal load, i.e. heat flow density pair
The integral of time, wall surface temperature, the heat flow density that active heat exchange amount, that is, active cooling is taken away.
This step is based on Engineering Algorithm and solves aircraft surface equation of heat balance acquisition Aerodynamic Heating.Aircraft surface balance side
Journey are as follows:
qaero=α (Tr-Tw)=ε σ Tw 4+qin
Q in formulaaeroFor Aerodynamic Heating heat flow density, α is aircraft surface convection transfer rate, TrFor gas recovery temperature, Tw
For aircraft surface temperature, ε is aircraft surface blackness, and σ is this special fence-Boltzmann constant.Method in the present invention is being counted
When calculating Aerodynamic Heating, it is added to the heat item q transmitted into aircraft cabinin, the as heat flow density taken away of active cooling.It is practical
In engineering, Active Cooling System is located at inner boundary, that is, cold end bottom of passive thermal protection system, the heat balance schematic diagram of system
As shown in Fig. 2, the item q that actively exchanges heatinAre as follows:
qin=h (Tb-Tc)
H, T in formulabAnd TcThe coefficient of heat transfer, passive heat respectively between Active Cooling System and passive thermal protection structure is anti-
The temperature of protecting system cold end allowable temperature and cooling working medium.The thermal protection system of active/passive coupling, actively the intensity of heat exchange is answered
The intensity with aircraft surface Aerodynamic Heating is corresponding, i.e., in wall surface temperature TwHigh region increases the intensity of actively heat exchange,
It is on the contrary then weaken heat transfer intensity.But in above formula, qinWith TwIt is unrelated, if substituting the above to surface balance equation, need with h for design
Parameter needs different h in the different region of Aerodynamic Heating intensity, therefore, so that design is extremely complex, is unable to complete successfully.
For this problem, current invention assumes that active cooling is located under the wall surface of passive thermal protection system, as shown in figure 3, at this point, actively
Exchange heat item qinIt is equal to:
qin=he(Tw-Tc)
In formula, qinWith TwIt is related, equivalent heat transfer coefficient h at this timeeIt can be used as design variable, for entire aircraft surface,
Identical heThe item that actively exchanges heat of Aerodynamic Heating intensity different zones can be simulated.
Equivalent heat transfer coefficient heRelationship between actual coefficient of heat transfer h is determined by following formula:
Based on above-mentioned model and it is assumed that using trajectory, aircraft configurations and equivalent heat transfer coefficient as input condition, Pu Lang is utilized
Special boundary layer theory, modified newton method and reference temperature method construct Engineering Algorithm program, can be obtained Aerodynamic Heating by calculating
Heat carries (integral of the heat flow density to the time), wall surface temperature and active cooling heat exchange amount.
Second step is the distribution for determining passive thermal protection concept.Metal heat-proof tile, the pottery developed in the present embodiment with NASA
Porcelain heat-proof tile and ceramic insulation felt construct thermal protection conceptual database, the density comprising each layer of thermal protection concept, maximum allowable temperature
Degree, radiant emissivity, thermal coefficient, specific heat and original depth.For region a certain for aircraft surface, to be obtained in the first step
The maximum wall surface temperature in the region is input condition under overall trajectory out, is aircraft wall region from thermal protection conceptual database
Domain selects thermal protection concept, and basic principle is that the maximum wall surface temperature in the region should be less than the allowable temperature of thermal protection concept.Needle
Suitable thermal protection concept is determined to aircraft surface all areas, the final distribution for obtaining the passive thermal protection concept of aircraft.
Third step is to determine that the scale of passive thermal protection system includes the appearance of thickness and quality and active thermal-protection system
Amount includes the mass flow of cooling working medium and the temperature rise of permission.It is established for the passive thermal protection concept that second step determines one-dimensional non-
Steady state heat transfer numerical model, as shown in figure 4, the aero-thermal load and active heat exchange amount with first step output are respectively outer, inner boundary
Condition carries out analysis of Heat Transfer, calculates the Temperature Distribution of the Temperature Distribution, that is, layers of material and boundary that obtain passive thermal protection concept.
Then, with inner boundary, that is, cold end set temperature TbAllowable temperature with layers of material is constraint, most with thermal protection system quality
Gently be target, in passive thermal protection concept thermal insulation layer with a thickness of design parameter, carry out structure optimization, final calculate obtains most
The thickness and quality of excellent passive thermal protection system.Specific heat with the active heat exchange amount of first step output and cooling working medium is defeated
Enter, calculates the temperature rise of the mass flow and permission that obtain working medium, the as capacity of active thermal-protection system.
This step is the capacity for calculating passive thermal protection system scale and active thermal-protection system.For passive thermal protection system
System, establishes the finite difference numerical model of one-dimensional and unsteady state heat transfer, with the cold end design temperature of thermal protection concept and each layer material
Expect that maximum allowable temperature is constraint, the scale for obtaining aircraft thermal protection concept, i.e. thickness are calculated by numerical analysis and optimization
And quality.Thermal protection concept is made of multilayer material, is based on its design feature, establishes the finite difference of one-dimensional multi-layer planar heat transfer
Model is as shown in Figure 4.Analysis node through-thickness is chosen, and the perpendicular bisector between adjacent node is interface, the area between adjacent interfaces
Domain is control volume representated by present node.Entirely zoning is discrete, needs according to first node location, rear interface position
The method set finally obtains the area of space of multiple subregion compositions.Further, since the mutual not phase of the thickness and physical property of layers of material
Together, need to carry out inside each layer it is careful discrete, and the size of unit by structure, temperature and calculation scale weigh choose, it is inside and outside
The borderline element thickness in boundary and gap is the half of other units.In addition, discrete region also needs to meet two conditions: heat
Node must be arranged by protecting in the inner and outer surfaces of concept, and the interface of two kinds of materials must be on the boundary of unit.
On the basis of discrete region, then governing equation can be carried out discrete.For control volume i (0 < i < N), thermal balance
Equation are as follows:
In formula, t is time, xiFor the coordinate position of node, ΔiFor the thickness of control volume i, ρ is the density of the layer material, c
For the specific heat of the layer material, k is the coefficient of heat conduction of the layer material.The formula is write into difference form, then is had:
In view of the Lamellar character of thermal protection structure, it can be assumed that in control volume, T, ρ, c be about x step variation,
Then it is believed that T, ρ, c value inside control unit are the value at cell node, therefore:
It is discrete with implied format progress, it takesImplicit Spline smoothing is made to the time, i.e., is used in the variation of t to t+ time Δt
Numerical representation method when t+ Δ t, obtains following formula:
Thermal coefficient between different material layer on interface is discontinuous, then the equivalent heat conductivity on interface is with harmonic average
Method, which calculates, to be obtained, to have:
Wherein WiFor thermal resistivity, value are as follows:
δ=1 when node i is located on boundary, remaining situation take 0, finally obtain the number of one-dimensional flat late heat transfer analysis model
It is worth discrete equation are as follows:
Rearrangement can obtain:
For outer boundary, third boundary condition, equation of heat balance are used herein are as follows:
Then the energy-balance equation of first control volume is writeable are as follows:
To obtain the discrete equation on outer boundary:
It is obtained after arrangement:
For inner boundary, with active heat exchange amount qinWrite as q for boundary condition for unified formatin=h (TN-Tc),
That is third boundary condition, so that it is writeable to obtain energy-balance equation on inner boundary control volume are as follows:
To which inner boundary discrete equation is writeable are as follows:
It is obtained after arrangement:
In conclusion obtaining considering the one-dimensional plate Transient Heat Transfer finite difference equations of active cooling:
By iterative solution, the Temperature Distribution of thermal protection concept structure can get, then with the allowable temperature of layers of material
And the design temperature of system cold end is constraint, is based on the minimum target of architecture quality using insulation thickness as design parameter
Sequential quadratic programming method carries out structure optimization, finally obtains optimal thermal protection system.The mathematics of the system optimization problem is retouched
It states are as follows:
s.t.Ti(x) < Timax, i=1,2 ..., N
TN(x) < Tb
In formula, n is the number of plies of thermal protection structure, and x is the thickness of each layer, ρi、Ti(x)、TbRespectively the density of layers of material,
Maximum temperature and system cold junction temperature, wherein TbThe input constraint designed for whole system.
, can be based on active heat exchange amount and working medium specific heat for active thermal-protection system, solving following formula can be obtained active thermal
The capacity of guard system, i.e. working medium flow and temperature rise:
qin=cmfΔT
In formula, c is the specific heat of cooling working medium, mfFor mass flow, △ T is working medium temperature rise.
4th step is the output active/passive thermal protection system of aircraft.The passive and active thermal protection obtained based on third step
After system, assess whether to meet Aircraft Conceptual Design demand, i.e., the quality and size of passive thermal protection system, active thermal protection
Whether working medium flow needed for system meets Aircraft Conceptual Design constraint, if it is satisfied, then current system can be used as aircraft
Active/passive thermal protection system；If not satisfied, then modifying active cooling parameter includes equivalent heat transfer coefficient and Temperature of Working, weight
Multiple step 1 carries out the design of a new round, such iteration is until finally meet Aircraft Conceptual Design demand to three.
Although the embodiments of the present invention has been shown and described above, it is to be understood that above-described embodiment is example
Property, it is not considered as limiting the invention, those skilled in the art are not departing from the principle of the present invention and objective
In the case where can make changes, modifications, alterations, and variations to the above described embodiments within the scope of the invention.
Claims (5)
1. a kind of active/passive thermal protection system coupling design method towards hypersonic aircraft totality, it is characterised in that: packet
Include following steps:
Step 1: being input ginseng with the 3-d modelling of hypersonic aircraft and grid division, task trajectory and active cooling parameter
Number is solved based on equation of heat balance utilizing works algorithm and obtains hypersonic aircraft Aerodynamic Heating, and aero-thermal load, wall surface temperature are exported
Degree and active heat exchange amount；The active cooling parameter includes equivalent heat transfer coefficient, cooling Temperature of Working；
Step 2: according to the wall surface temperature obtained in step 1, determining maximum wall surface temperature of the analyzed area under overall trajectory；To divide
Analysing maximum wall surface temperature of the region under overall trajectory is input condition, is superb from the thermal protection conceptual database established in advance
The analyzed area of velocity of sound aircraft wall surface selects thermal protection concept, it is desirable that maximum wall surface temperature of the analyzed area under overall trajectory
Less than the allowable temperature of selected thermal protection concept, and then obtain the passive thermal protection concept point of entire hypersonic aircraft
Cloth；
Step 3: establishing one-dimensional and unsteady state heat transfer numerical model for the passive thermal protection concept that step 2 determines, and obtained with step 1
The aero-thermal load and active heat exchange amount arrived is respectively outer, internal boundary condition progress analysis of Heat Transfer, calculates passive thermal protection concept
Temperature Distribution；It is then constraint with the set temperature of inner boundary and the allowable temperature of passive thermal protection concept layers of material, with quilt
Dynamic thermal protection system quality is most gently target, in passive thermal protection concept thermal insulation layer with a thickness of design parameter, carry out structure
Optimization calculates the thickness and quality for obtaining optimal passive thermal protection system；It is obtained with cooling down specific heat and the step 1 of working medium
Active heat exchange amount is input, calculates the temperature rise of the mass flow and permission that obtain cooling working medium,
Step 4: cooling needed for the quality and thickness, active thermal-protection system of the passive thermal protection system that judgment step 3 obtains
Whether working medium mass flow meets hypersonic aircraft overall design constraints, if it is satisfied, then current system can be used as flight
The active/passive thermal protection system of device；If not satisfied, then modifying active cooling parameter, step 1 is repeated to 3, carries out setting for a new round
Meter, such iteration is until finally meet hypersonic aircraft overall design constraints.
2. a kind of active/passive thermal protection system Coupling Design towards hypersonic aircraft totality according to claim 1
Method, it is characterised in that: task trajectory includes flight time, height, speed, the angle of attack and yaw angle in step 1.
3. a kind of active/passive thermal protection system Coupling Design towards hypersonic aircraft totality according to claim 1
Method, it is characterised in that: equation of heat balance described in step 1 be Aerodynamic Heating amount be equal to radiation dissipation amount and active heat exchange amount it
With；The active heat exchange amount is equal to wall surface temperature and cooling working medium temperature difference multiplied by equivalent heat transfer coefficient.
4. a kind of active/passive thermal protection system Coupling Design towards hypersonic aircraft totality according to claim 1
Method, it is characterised in that: the Engineering Algorithm used in step 1 is based on prandtl boundary layer theory, is regarded as ideal gas outside boundary layer
Body has viscous gas in boundary layer, by boundary layer outside calculating obtain aerodynamic parameter outside boundary layer, utilize reference temperature later
Method calculates the Aerodynamic Heating obtained in boundary layer.
5. a kind of active/passive thermal protection system Coupling Design towards hypersonic aircraft totality according to claim 1
Method, it is characterised in that: in step 2 thermal protection conceptual database include the density of each layer of thermal protection concept, maximum allowable temperature,
Radiant emissivity, thermal coefficient, specific heat and original depth.
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114491791A (en) * | 2021-12-28 | 2022-05-13 | 中国航天空气动力技术研究院 | Design method of controllable ablation heat-proof structure for aircraft |
| CN114647260A (en) * | 2022-03-19 | 2022-06-21 | 西北工业大学 | Active thermal control method of aircraft servo motor based on variable-density structure micro-channel heat sink performance regulation and control |
| CN114839864A (en) * | 2022-07-04 | 2022-08-02 | 中国飞机强度研究所 | Radiation interference decoupling control method of aerospace plane heat intensity test control system |
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| CN114647260B (en) * | 2022-03-19 | 2023-08-29 | 西北工业大学 | Aircraft servo motor active heat control method based on variable density structure micro-channel heat sink performance regulation and control |
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| CN114839864B (en) * | 2022-07-04 | 2022-09-13 | 中国飞机强度研究所 | Radiation interference decoupling control method of aerospace plane heat intensity test control system |
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